The invention relates to inertial reference systems (IRS) and to attitude and heading reference systems (AHRS), in particular those based on rate gyros using microelectromechanical systems (MEMS) technology.
Controlling a moving body (e.g. an aircraft) requires inertial measurements to be taken relating to the six degrees of freedom of the moving body. As a general rule, these are usually firstly measurements of the three components of the angular velocity vector, and secondly of the three components of the angular acceleration vector.
Historically, angular measurements were initially made by means of free gyros, and subsequently they have been made by means of rate gyros that measure the angular velocity (rotation) components of the carrier directly.
Rate gyros include in particular so-called “strap-down” gyros (i.e. their axes of rotation are constrained to remain parallel to the axis of the carrier, with the applied force being proportional to angular velocity), laser gyros, optical fiber laser gyros, and resonating structure gyros.
In a resonating structure gyro, a mechanical resonator (such as a tuning fork) is caused to vibrate and its oscillations are sustained, with the movements thereof perpendicular to the excitation plane being measured. Coriolis forces tend to keep the vibration plane fixed in an inertial frame of reference, so such perpendicular components appear only in the presence of angular velocity and they are proportional to the amplitude thereof. That type of resonator can be miniaturized down to a scale of a MEMS made of silicon and located in an integrated circuit, thereby making it possible to fabricate a gyro at low cost.
Nevertheless, in such a gyro, since the resonating mass is extremely small, measurement noise is high. In a precision inertial unit, use is generally made of laser gyros having an intrinsic noise level that is of the order of 100th the noise level of a microsensor (of the MEMS type). It is known to incorporate angular accelerometers in a strap-down inertial unit in order to attempt to correct its deterministic errors (improperly referred to as “high frequency noise”) as constituted by the cone and sculling effects that appear during dynamic stages of flight and in the event of computations being performed at too slow a rate or of the gyros having too narrow a passband. The amplitude of these errors is troublesome in navigation grade inertial units, but not for autopilot sensors, particularly since there is no longer a computation rate limitation given the power of modern computers.
These navigation grade gyros are laser rings of large size or possibly fiber optic gyros (FOGS), likewise of large size. Navigation applications are not accessible to MEMS inertial sensors. Rate gyros are essential sensors for an aircraft autopilot (below “AP”). It is possible to model a system including an aircraft 20 and it AP as shown in FIG. 2.
The main purpose of an AP is to stabilize the aircraft when faced with disturbances caused by turbulence in the mass of air. One way of modeling the effect of such turbulence is to represent it as a term 21 that is added to the movements of the flight control actuators 22.
The diagram of FIG. 2 serves to establish the following transfer function (where w is the actual angular velocity, P is the disturbance, B is the noise of the gyro 24, C is the gain of a correcting filter 23 for correcting gyro measurements, and where the transfer functions of the aircraft, of the actuator, and of the gyro are taken to be unity):
  ω  =            (                                                  1                              1                +                C                                                                        C                              1                +                C                                                        )        ⁢          (                                    P                                                B                              )      
Given the simplifications that are adopted, the corrector reduces to an integrator:
  C  =                    2        ⁢        π        ⁢                                  ⁢                  f          0                    p        =          1              τ        ⁢                                  ⁢        p            where f0 is the closed loop resonant frequency of the airplane with its autopilot. The transfer function then takes the form:
  ω  =            (                                                                  τ                ⁢                                                                  ⁢                p                                            1                +                                  τ                  ⁢                                                                          ⁢                  p                                                                                        1                              1                +                                  τ                  ⁢                                                                          ⁢                  p                                                                        )        ⁢          (                                    P                                                B                              )      
It can be seen that the system is complementary. It applies a highpass filter to the disturbances and a lowpass filter to the angular velocity measurement noise, using the same cutoff frequency. If the resonant frequency (i.e. the open loop gain) is increased to reject disturbances, then the bandwidth of the lowpass filter is increased in equal manner, thereby transmitting the sensor noise to the entire airplane.
When developing a helicopter AP fitted with FOGs that nevertheless present low measurement noise, the limiting factor on increasing the gain of the corrector is measurement noise, which is manifested by the appearance of broadband vibration felt by the crew. It is therefore measurement noise, even in high quality gyros, that limits the overall performance of the loop. Most present autopilots take advantage of the low noise of FOGs, in spite of their expense.
The graph of FIG. 1 shows variations in spectral power density (PSD)—in degrees per second per square root of hertz (°/s/√Hz)—of the angular measurement noise respectively of a FOG, of a closed loop (CL) type MEMS, and of an open loop (OL) type MEMS, plotted up the ordinate as a function of frequency, plotted along the abscissa.
Given that the frequency band for an autopilot extends well beyond 1 hertz (Hz) (where the typical passband of a helicopter AP gyro is 10 Hz), and given the noise level of a FOG is the limiting criterion on improving an AP in terms of its response to turbulence, a MEMS gyro, even one of the closed loop type, presents a noise level that is excessive.
Furthermore, an “f” noise profile (i.e. a profile that increases in proportion to frequency) makes gain adjustment more sensitive: unlike a FOG in which noise amplitude increases with the square root of the passband, the noise level transmitted by the MEMS increases directly with frequency.